The present invention generally relates to attitude control and rate sensing of a vehicle, such as a spacecraft, and more particularly relates to a control moment gyroscope for providing inertial control torque and attitude rate sensing.
Control moment gyroscopes (CMGs) are commonly used to provide precise directional steering control torque for a variety of vehicles, such as spacecraft satellites to maneuver the spacecraft. A control moment gyroscope is mounted to the spacecraft and typically includes a motor driven rotor spun about a rotor axis, a gimbal, a gimbal torque motor to rotate the gimbal about a gimbal axis, and a control system. The rotor is supported on the gimbal and is rotated about the gimbal axis which is generally perpendicular to the rotor axis. One example of a control moment gyroscope is disclosed in U.S. Pat. No. 5,386,738, which is hereby incorporated herein by reference. The conventional control moment gyroscope is operated such that the rotor is spun about its rotor axis at a predetermined rate to generate angular momentum. In order to induce a torque on the spacecraft, the gimbal torque motor rotates the gimbal and thus the spinning rotor about the gimbal axis. Rotation of the stored angular momentum vector produces a significant output torque which is perpendicular to the rotor axis and the gimbal axis. The output torque, in turn, is transferred directly to the spacecraft. In a typical spacecraft application, an array of at least three control moment gyroscopes are often used to produce control torque in any direction.
In order to determine the amount of output torque to induce with each control moment gyroscope, the control system receives commands and sensed parameters of the spacecraft including spacecraft angular attitude rates for yaw, pitch, and roll. The sensed attitude rates are monitored to generate control signals to operate the control moment gyroscope(s) to maintain stability and dynamic control of the spacecraft. To measure the inertial spacecraft attitude rates, the spacecraft generally employs separate gyroscopic units, referred to as rate sensing gyros. Each rate sensing gyro has a rotor, gimbal, and motor assemblies or other sensing means which add additional mass, weight, and cost to the spacecraft. Also, backup rate sensing gyros are typically mounted on the spacecraft and are employed in the event of a failure of a primary rate sensing gyro. The use of additional backup rate sensing gyros further adds to the overall mass, weight, and cost. Further, in instances where the control moment gyroscope is sufficiently vibrationally isolated from the spacecraft, the rate sensing gyro and control moment gyroscope may not be co-located which made lead to potential servo control problems, particularly with very high control bandwidth attitude control systems.
Accordingly, it is therefore desirable to reduce the weight and cost for control actuation and inertial measurement on a vehicle. In particular, it is desirable to integrate functions of the control moment gyroscope and attitude rate sensing gyro to thereby minimize the weight and cost added to a satellite spacecraft and co-locate the attitude control and sensing in some cases.
In accordance with the teachings of the present invention, an inertial control and measurement system and method are provided for generating inertial control torque and measuring angular rate of a vehicle. The system employs a control moment gyroscope having a rotor adapted to be spun by a rotor motor, a gimbal, a gimbal support assembly adapted to be attached to a vehicle for allowing rotation of the gimbal about a gimbal axis, and a gimbal motor for rotating the gimbal about the gimbal axis to induce torque. The system further includes a controller for controlling the gimbal motor to generate the control torque. The controller further determines an attitude angular rate of the vehicle as a function of a determined torque and angular acceleration of the gimbal.
According to another aspect of the present invention, a control moment gyroscope is provided for applying torque to a vehicle and determining an angular attitude rate of the vehicle. The control moment gyroscope includes a rotor adapted to be rotated about a rotor axis, a gimbal supported by an assembly attached to a vehicle, and a gimbal motor for rotating the gimbal about a gimbal axis. The control moment gyroscope further includes an acceleration determining device for determining an acceleration of the gimbal about the gimbal axis, and a torque determining device for determining torque applied to the gimbal shaft. A controller controls the gimbal motor to generate a control torque, and further determines an angular rate of the vehicle as a function of the determined torque and a gimbal acceleration. Accordingly, the control moment gyroscope provides integrated inertial control and angular rate measurement.